The assignee of the present invention manufactures and deploys spacecraft for, inter alia, communications and broadcast services from geostationary orbit. During launch, such spacecraft are enclosed within a launch vehicle payload fairing that experiences depressurization from an initial pressure of nominally one atmosphere (approximately 14.7 PSIA) to a near vacuum condition within a time period of about two minutes. To safely accommodate this pressure change, the spacecraft design must include provisions for safely venting of air from interior volumes of the spacecraft and spacecraft components into the launch vehicle payload fairing.
Large structural components of a spacecraft include spacecraft equipment and solar array panels and structural panels that may be in the range of 50 square feet surface area, or greater. Referring to FIG. 1, an exploded isometric view of an example structural panel 100 is illustrated. Panel 100 includes a honeycomb core 110 sandwiched between panel faceskins 120. The panel faceskins 120 may be formed from aluminum or a carbon composite material, for example, and have a cell wall thickness, typically, of less than 0.01 inches. The panel faceskins 120 may be adhered to a honeycomb core 110 by epoxy adhesive or other adhesive bond, for example.
Referring now to views A-A and B-B of FIG. 1, each cell in the honeycomb core is intended to be vented (by slitting or perforating, for example) to permitted air to escape during launch as the spacecraft leaves the earth's surface and experiences a depressurization from approximately one atmosphere of pressure to the vacuum of space.
In practice, however, it has been found that some cells of an as-fabricated honeycomb panel, which may typically include several thousand cells, may exhibit manufacturing defects as a result of which the defective cells fail to comply with the design intent of providing safe venting means. Such manufacturing flaws are difficult to completely prevent and may be difficult to detect by conventional inspection or nondestructive test techniques. A consequence of such undetected flaws can include explosive rupture of the panel, and resulting damage to spacecraft functional systems.
The bond strength between panel faceskin and core may vary from panel to panel and within a given flight panel. To mitigate the risk of damage in the event that normal venting of the honeycomb core during launch ascent is not achieved, it is desirable that the bond strength be sufficient to withstand a pressure differential of one atmosphere or more. Bond strength between panel faceskin and core of a flight panel is ordinarily estimated by way of destructively testing coupon samples that are co-manufactured with the flight panel. However, such coupon testing has been problematic because coupon bond strength may not accurately correlate with bond strength of the flight panel, particularly the weakest point of the flight panel.
As a result, an improved approach to bond strength testing of flight honeycomb core panels is desirable.